The present invention relates to gas turbine engines, and in particular, to thermal conditioning of bore cavities in gas turbine engines. Gas turbine engines typically operate with very high temperatures in compressor, combustor, and turbine sections. Different components in a gas turbine engine are subjected to different temperatures depending on component location. Additionally, a single component can be subjected to different temperatures at different parts of that component. Moreover, a component can be exposed to different temperatures during different operating conditions of the gas turbine engine.
For example, when hot gas flows through a compressor section during engine idle, disks that form a rotor in the compressor section can have a relatively hot rim and a relatively cool bore and web. When the gas turbine engine is sped up, for example, prior to take-off of an aircraft for a propulsion engine, the hot gas can increase in temperature. This can further increase the temperature of rims of the disks. While the bore and web of the disks may eventually increase in temperature as well, this can happen at a rate slower than that of the rim of the disk. Because metal disks typically expand when heated, this can result in a situation where disk rims expand more quickly than disk webs and bores, this creating undesirable stresses in the webs and bores. The opposite effect can happen during aircraft descent, where disk rims cool more quickly than disk bores and webs. This can also result in undesirable stresses in the webs and bores, which can lead to damaged disks after a number of cycles.